Higher operating temperatures for gas turbine engines are continuously sought in order to increase their efficiency. However, as operating temperatures increase, the high temperature durability of the components of the engine must correspondingly increase. Significant advances in high temperature capabilities have been achieved through formulation of iron, nickel and cobalt-based superalloys, though components formed from such alloys often cannot withstand long service exposures if located in certain sections of a gas turbine engine, such as the turbine, combustor or augmentor.
A common solution is to provide internal cooling of turbine, combustor and augmentor components, at times in combination with a thermal barrier coating. Airfoils of gas turbine engine blades and vanes often require a complex cooling scheme in which cooling air flows through cooling channels within the airfoil and is then discharged through carefully configured cooling holes at the airfoil surface. Convection cooling occurs within the airfoil from heat transfer to the cooling air as it flows through the cooling channels. In addition, fine internal orifices can be provided to direct cooling air flow directly against inner surfaces of the airfoil to achieve what is referred to as impingement cooling, while film cooling is often accomplished at the airfoil surface by configuring the cooling holes to discharge the cooling air flow across the airfoil surface so that the surface is protected from direct contact with the surrounding hot gases within the engine.
In the past, cooling channels have typically been integrally formed with the airfoil casting using relatively complicated cores and casting techniques. More recently, U.S. Pat. Nos. 5,626,462 and 5,640,767, both to Jackson et al. and commonly assigned with the present invention, teach a method of forming a double-walled airfoil by depositing an airfoil skin over a separately-formed inner support wall (e.g., a spar) having surface grooves filled with a sacrificial material. After the airfoil skin is formed, preferably by deposition methods such as plasma spraying and electron-beam physical vapor deposition (EBPVD), the sacrificial material is removed to yield a double-walled airfoil with cooling channels that circulate cooling air against the interior surface of the airfoil skin.
In Jackson et al., the sacrificial material is a braze alloy deposited in excess amounts in a spar groove, with the excess being removed by machining or another suitable technique so that the surface of the braze alloy is flush with the surrounding surface of the spar. The sacrificial material is then removed after deposition of the airfoil skin by melting/extraction, chemical etching, pyrolysis or another suitable method. Though braze alloys have been successful in the process disclosed by Jackson et al., efforts have continued to develop other materials that better meet the requirements described previously. Notably, sacrificial materials proposed for other applications have been tried without success. For example, a combination of K.sub.2 SO.sub.4 and Na.sub.2 AlO.sub.3 was experimented with as a sacrificial backfill material, but found to be corrosive and severely attacked a spar formed of Rene N5, a General Electric nickel-based superalloy having a nominal composition, in weight percent, of Ni-7.5Co-7.0Cr-6.5Ta-6.2Al-5.0W-3.0Re-1.5Mo-0.15Hf-0.05C-0.004B-0.01Y. Other known sacrificial materials, including those disclosed in U.S. Pat. No. 4,956,037 to Vivaldi and U.S. Pat. No. 5,249,357 to Holmes et al., are unable to withstand high-temperature deposition processes such as EBPVD.
In view of the above, it would be desirable if improved sacrificial materials and processes were available that could ensure that all of the aforementioned requirements of the sacrificial material were adequately met to produce an air-cooled component with a deposited skin that is substantially free of surface defects.